Apparatus for gyroscopic ejection of shuttle launched spacecraft

ABSTRACT

A spacecraft specifically adapted for launch from a reusable launch vehicle such as the space shuttle by means of a cradle having locking and ejection mechanisms mounted therein. The cradle fastens into the payload bay of the space shuttle and returns therewith for reuse in subsequent launches. The spacecraft mounts at three points to the cradle, and the cradle mounts at three points to the shuttle such that a plane through the attachment points passes through the roll axis of the spacecraft at approximately the center of mass thereof. The cradle utilizes the truss structure of the spacecraft to produce the required stiffness by providing a structural tie between the two ends and the bottom of the cradle. At launch, the spacecraft is ejected with both linear and angular momentum, the spin providing gyroscopic stability. The locking mechanisms in the cradle can be remotely controlled to relock the spacecraft to the cradle in the event of an unsuccessful deployment attempt. The spacecraft includes a safety circuit employing acceleration sensing switches which sense spin-up of the spacecraft and prevent premature ignition of the perigee boost motor. The spacecraft has imbedded within its envelope a solid-propellant perigee boost motor surrounded by a liquid-propellant apogee motor. By employing apogee and perigee propulsion stages internal to the spacecraft, the storage length in the space shuttle is minimized, and the geometry and mass characteristics of the spacecraft make for a stable spinning vehicle during both the perigee and apogee boost phases.

BACKGROUND OF THE INVENTION

Geostationary satellite systems have already demonstrated the ability toperform many communication, meteorological, and scientific missions whenlaunched by expendable launch vehicles. The NASA Space TransportationSystem (STS), the Space Shuttle, offers the opportunity for asignificant improvement in the performance and cost of satellitesdesigned to take advantage of its capabilities. It therefore can makecurrent applications more profitable and new uses economicallyattractive.

All satellite systems require launch vehicles, and all current launchvehicles are expended after delivering their spacecraft payload intoorbit. The NASA Space Shuttle by contrast, introduces a new concept ofbeing recoverable and reusable. Tests have proved that the Space Shuttlecan be piloted like an airplane after re-entering the atmosphere fromspace.

All geostationary satellites so far have been launched by Thor Delta,Atlas Agena, Atlas Centaur, or Titan IIIC launch vehicles. Now, however,the spacecraft designer has the choice of six launch vehicles: ThorDelta (2914, 3914, 3910 PAM), N-Rocket, Atlas Centaur, Ariane, TitanIIIC, and the Space Transportation System. The N-Rocket is beingdeveloped by Japan's NASDA (National Space Development Agency) and theAriane by Europe's ESA (European Space Agency). The United States plansto phase out the Thor Delta, Atlas Centaur, and Titan IIIC as the STSbecomes operational in 1980. The returnable and reusable Space Shuttleoffers the challenge and opportunity to geostationary spacecraftdesigners to make the best use of it.

The Space Shuttle will orbit Earth at a nominal 160 n.mi. with an orbitinclination of 28.6 deg when launched due east from Florida. Ageostationary satellite must orbit at approximately 19,300 nauticalmiles north of the equator. The STS therefore needs an upper stage tolaunch geostationary satellites. The upper stage requirements areoptimally satisfied by two propulsion impulses. At the time of anequatorial crossing, the first impulse imparts a velocity increment ofapproximately 8000 fps at the perigee of elliptical transfer orbit. Atan appropriate apogee of the transfer orbit the second impulse imparts avelocity increment of 6000 fps, both circularizing the orbit andremoving the inclination.

The central challenge in using the STS for the launch of geostationarysatellites lies in finding the combination of upper stage and satellitegeometry and functions that minimizes overall mission cost.

Upper-Stage Alternatives: The first upper-stage concepts considered forthe STS completely separated orbit-injection functions from subsequentorbit-control requirements. These bulky and expensive "stand alone"upper stages obscure the basic economic advantages of the STS.

It soon became apparent that the STS could take advantage of thegeostationary orbit-injection scheme pioneered by Syncom III in anhistoric Delta launch in 1964. This early capability was achieved byincorporating the apogee-boost motor within the satellite, thuspermitting separation from the Delta at perigee injection. The perigeeboost itself was provided primarily by the Delta's unguided, spinningupper stage. Since the apogee boost does not parallel the perigee boost,it was necessary to reorient the spacecraft spin axis beforeapogee-motor firing. This was done by the control and attitude sensingsystem required for Syncom's operational mission. The use of thesatellite's telemetry and command system and communication repeaters todetermine the transfer-orbit parameters via ground tracking permittedthe selection of an apogee-motor firing time and attitude that minimizedthe effects of transfer-orbit injection errors. The subsequent launch ofsome 50 geostationary satellites by the Delta and Atlas Centaur boostersbrought refinements but no basic changes to this technique.

NASA adopted the Spinning Solid Upper Stage (SSUS) as its preferredmethod of launching via the STS geostationary spacecraft previouslydesigned for expendable launch vehicles. These upper stages are nowknown as SSUS-D for Delta and SSUS-A for Atlas Centaur replacements. TheUSAF has elected to retain the independent upper stage concept used inthe Titan IIIC for its STS launches of spacecraft to high-energy orbits.It will use a pair of guided and controlled solid-propellant rocketstages known as the IUS (for Interim Upper Stage).

The STS/SSUS does offer a lower launch cost than the expendable launchvehicles for organizations unwilling or unable to depend exclusively onSpace Shuttle operational availability. The accommodation of transitionspacecraft in the Space Shuttle, while retaining their ability to belaunched on expendable launch vehicles, has proven to be achievable. Inthe case of new spacecraft, the dual capability can readily beincorporated in the initial design.

Even for transition spacecraft, the STS reduces launch costssignificantly. The STS economics clearly support the desirability ofincorporating dual capability into transition spacecraft, permittingSpace Shuttle launch while maintaining expendable-launch-vehicle backup.

The transition spacecraft, attractive as they are, do not representoptimum designs for the Space Shuttle, for two reasons. First therestriction on diameter imposed by the expendable launch vehicles makesthem longer than otherwise necessary. The SSUS-D and SSUS-Aconfigurations, for example, require a quarter, and a half,respectively, of the Space Shuttle payload bay volume, but use a muchsmaller fraction of the available weight. The dual compatibility thuslevies a penalty of higher than necessary launch cost of the Spaceshuttle. As another disadvantage, such dual-launch spacecraft do nothave as much space for mounting antennas, cameras, and scientificinstruments as would one designed only for Shuttle launch.

This invention relates to the apparatus and method for payloaddeployment from a space shuttle and more particularly to an apparatusfor the gyroscopic ejection of a spacecraft from the shuttle.

SUMMARY OF THE INVENTION

A primary object of the subject invention is to provide a new andimproved method and apparatus for ejecting a payload from a spaceshuttle-type vehicle.

A more specific object of the present invention is to provide a costeffective and reliable apparatus for gyroscopically ejecting a payloadfrom a space vehicle.

One embodiment of the present invention which is particularly suited forlaunching a spacecraft from a space shuttle, comprises a U-shaped cradlehaving an injection spring located on one side thereof for pushingagainst a small trunnion that protrudes from one side of the spacecraft.A pivot point is formed on the opposite side of the spacecraft whichalso has a protruding trunnion that rests on a mating surface formed inthe cradle. Both the spring and pivot points lie in a plane normal tothe spin axis of the spacecraft and, ideally, a plane also passingthough the center of mass of the spacecraft. Release of the ejectionspring causes the spacecraft to rotate about the pivot point creating atranslation and rotation about the spacecraft center of mass. After theejection force ceases, the spacecraft free body motion is a rollingmotion up a virtual (imaginary) ramp thus maintaining the impartedlinear and angular momentum. The separation velocity and rotation speedof the spacecraft depends on its inertia characteristics, diametricdimensions, ejection force and stroke length.

In accordance with the deployment method of the invention, thespacecraft would be stored in the bay of the shuttle with the spacecraftspin axis aligned parallel to the roll axis of the shuttle whereby thespacecraft attitude is set by the attitude of the shuttle and maintainedduring ejection by simultaneously imparting both linear and angularmomentum to the spacecraft. As a consequence, no active control isrequired prior to firing the perigee propulsion stage of the spacecraft.A further advantage of the invention is that the ejection of thespacecraft is accomplished by a single mechanization which is comprisedof an ejection spring on one side and a pivot point on the other sidewith both of thes points preferably lying approximately in the plane ofthe center of mass of the spacecraft. Also, the injection mechanismwhich imparts linear and angular momentum to the spacecraft provides aseparation velocity as well as gyroscopic stability. The advantages ofthe mechanization in accordance with the subject invention include:

spacecraft attitude can be maintained without the need for activecontrol;

gyroscopic stability also helps to maintain adequate clearance as thespacecraft leaves the shuttle bay;

spacecraft spin insures propellant feed at thrustors which are to beused subsequently for additional spinup; and single pushoff and pivotpoints provide a simple mechanization.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention, including its various objects, features, and advantages,may be more readily understood with reference to the following detaileddescription of an embodiment thereof, taken in conjunction with theaccompanying drawings, wherein like reference numerals designate likestructural elements, and in which

FIG. 1 is a drawing of a space shuttle showing the payload bay doorsopen, and having a portion broken away to show a spacecraft in asupporting cradle mounted therein;

FIGS. 2a and 2b are diagrams illustrating the ejection of the spacecraftfrom the payload bay of the space shuttle;

FIGS. 3a, 3b and 3c are diagrams which depict the motion of thespacecraft during ejection from the shuttle;

FIG. 4 is a diagram showing the relationship of the parking orbit of thespace shuttle to the transfer orbit and synchronous orbit of thespacecraft;

FIG. 5 is a perspective view of a portion of the payload bay of thespace shuttle with the cradle mounted therein;

FIG. 6 is a side elevation view of one of the fittings which attachesthe cradle to the payload bay of the space shuttle;

FIG. 7 is a perspective view of two segments of one side of the cradle,partly broken away, showing bulkhead, skin and sides;

FIG. 8 is an enlarged perspective view of one of the bulkheads of thecradle;

FIG. 9 is a schematic diagram of the spacecraft illustrating how itmates to the cradle;

FIG. 10 shows a spacecraft trunnion ball in its swivel blockillustrating how the ball is locked into the cradle by a locking bolt;

FIG. 11 is a schematic diagram illustrating the forces on the threesatellite trunnion balls at the time of ejection;

FIG. 12 is a perspective view of the locking mechanism which drives thelocking bolt that locks two of the satellite trunnion balls to thecradle;

FIGS. 13, 14, 15 and 16 show the latching or locking mechanism used atthe point where the ejection spring applies its force. FIGS. 13 and 14are perspective views, while FIGS. 15 and 16 are side elevations takenat 90 degrees with respect to each other. FIGS. 13, 15 and 16 show themechanism in the locked position, while FIG. 14 shows the mechanism inthe released or unlocked position;

FIG. 17 shows the locking mechanism for the ejection spring mounted inthe end of the cradle and protruding into the slot provided for thespacecraft trunnion ball;

FIG. 18 shows the basic skeletal structure of the spacecraftillustrating how the truss structure of the spacecraft provides astructural tie between the two ends of the cradle and the bottom or keelpoint;

FIG. 19 is a side elevation of the spacecraft showing the truss framestructure in the axial direction;

FIG. 20 is a top plan view of the spacecraft showing the framestructure, solar panel, propellant tanks and perigee motor;

FIG. 21 is a perspective view of the spacecraft illustrating theorientation of the axes of the acceleration sensing g-switches; and

FIG. 22 is a circuit diagram showing the connection of the separationswitches and the g-switches.

DETAILED DESCRIPTION

Referring to FIG. 1 of the drawings, there is shown a space shuttle 50developed by the National Aeronautics and Space Administration. Thespace shuttle 50 is shown with its payload bay doors open, and a portionis broken away to show a spacecraft 51 mounted in the payload bay bymeans of a cradle 52. The payload bay of the space shuttle 50 is 15 feetin diameter, 60 feet long, and carries 65,000 pounds. It has beencalculated that a railroad box car can be loaded in the payload bay, andthe doors can be closed.

The spacecraft 51 is designed to take advantage of the capabilities ofthe space shuttle 50, and in accordance with this optimized design, thelength of the spacecraft 51 is nominally 10 feet and the diameter ismade as large as possible, namely 14 feet. The cradle 52 mounts in thepayload bay of the space shuttle 50 and simplifies the mechanicalattachment of the spacecraft 51 to the space shuttle 50, as well assimplifying ground handling procedures. In addition, the cradle 52provides the necessary locking and ejection mechanisms.

The spacecraft 51 is stowed in the bay of the space shuttle 50 with thespin axis parallel to the roll axis of the shuttle 50. Thus, theattitude of the spacecraft 51 is set by the attitude of the shuttle 50and is maintained during ejection by simultaneously imparting bothlinear and angular momentum to the spacecraft 51. The spin providesgyroscopic stabilization. FIG. 2a shows the spacecraft 51 beforeejection. The cradle 52 provides a pivot point 53 at one side of thespacecraft 51 and provides an ejection spring 54 at the other side ofthe spacecraft 51. Both the spring 54 and the pivot point 53 lie in aplane normal to the spin axis and ideally, also passing through thecenter of mass of the spacecraft 51. Release of the ejection spring 54causes the spacecraft 51 to rotate about the pivot point 53 creating atranslation and rotation about the center of mass of the spacecraft 51.After the ejection force ceases, the free body motion of the spacecraft51 is a rolling motion up a virtual or imaginary ramp 55, thusmaintaining the imparted linear and angular momentum. The separationvelocity and rotational speed depends on the inertia characteristics anddiametric dimensions of the spacecraft 51, and on the ejection force andstroke length.

FIG. 2b shows the spacecraft 51 during ejection, the series of lines 56indicating the motion of the spacecraft 51. The motion of the spacecraft51 during ejection may be even more clearly seen from FIG. 3. FIG. 3ashows the spacecraft 51 before ejection, FIG. 3b shows the spacecraft 51during ejection and FIG. 3c shows the spacecraft 51 after ejection. Thearrows indicate the motion of the spacecraft 51, which motion may be onthe order of 2-foot per second linear translation and two revolutionsper minute.

Referring now to FIG. 4, the space shuttle 50 is designed to orbit theearth at a nominal 160 nautical miles with an orbit inclination of 28.6degrees when launched due East from Kennedy Space Center at CapeCanaveral, Fla. This orbit is termed the parking orbit 57. To place thespacecraft 51 in geosynchronous orbit, a perigee kick motor is fired atthe perigee 58 of a transfer orbit 60, and an apogee kick motor is firedat the apogee 61 of the transfer orbit 60, such that the spacecraft 51is placed into the 19,300 nautical mile synchronous orbit 62.

The advantage of this type of ejection is that no active control of thespacecraft 51 is required prior to firing the perigee propulsion stage.Also, the ejection is accomplished by a simple mechanization which willbe described more fully hereinafter. The spin of the spacecraft 51provides gyroscopic stability although after the spacecraft 51 has beenejected, additional spin is imparted to it by spin-up jets. Thegyroscopic stability helps to maintain adequate clearance as thespacecraft 51 leaves the payload bay of the shuttle 50. The spin of thespacecraft 51 insures propellant feed at the thrusters which are usedsubsequently for additional spin up. The U-shaped cradle 52 remains withthe shuttle 50 after ejection of the spacecraft 51 and may be reused forsubsequent launches.

FIG. 5 shows the cradle 52 mounted in the cargo bay of the space shuttle50. The cradle 52 is provided with three trunnions 65, 66, 67, (see alsoFIG. 9) one at each side and one at the bottom. These trunnions 65, 66,67 mate with fittings 70, 71, 72 located on the shuttle 50. Two of thetrunnions 65, 67 are connected to longerons on each side of the shuttle50, and the bottom trunnion 66 is connected to the keel of the shuttle50. The fittings 70, 71, 72 provide a moment-free attachment and also aslip-joint in the radial directions. As a result, the U-shaped cradle 52is highly flexible when radial loads are applied at the longeronlocations.

FIG. 6 shows in more detail one of the trunnions 65 fastened to theshuttle 50 by one of the fittings 70. The three-point attachment betweenthe cradle 52 and the shuttle 50 provides a statically determinant(non-redundant) connection, thus eliminating loads into the spacecraft51 induced by installation tolerances and by distortions of the shuttle50 caused by either mechanical or thermal loading on the shuttle 50.These three attachment points lie in the same plane as the attachmentpoints for the spacecraft 51. As will be seen in FIG. 5, slots 75, 76,77 are provided in the cradle 52 for attachment of the spacecraft 51.

The three-point attachment between the spacecraft 51 and the cradle 52represents the minimum number of release points and thus the fewestnumber of release mechanisms which must be actuated before ejection canoccur. The plane formed by the attachment points passes through thelongitudinal (or roll) axis of the spacecraft 51 at approximately thelocation of the center of mass of the spacecraft 51. This leads tofavorable reductions in spacecraft loads and thus structural weight.This location is selected to reduce launch loads by eliminating thecantilever effect and to permit the cradle 52 to be the reactionplatform for the push-off mechanism. The push off is accomplished by theejection spring 54 with the spring force having zero longitudinal offsetfrom the center of mass of the spacecraft 51.

The cradle 52 utilizes the structure of the spacecraft 51 to provide therequired stiffness. In this way, it is possible to support a largediameter spacecraft 51 using a narrow u-shaped cradle 52. By utilizingthe spacecraft 51 to augment the stiffness of cradle 52, the cradle 52can be made sufficiently narrow to fit within the space between thespacecraft 51 and the 15 foot payload bay envelope. Thus, the weight ofthe cradle 52 is much less than if it had to be self-sustaining, andconsequently more weight may be allotted to the spacecraft 51 for thesame launch weight. The structure of the spacecraft 51 provides astructural tie between the two ends of the U-shaped cradle 52 and thekeel, and this tie greatly increases the overall stiffness of theintegrated system. In order to provide the required stiffness, thestructure of the spacecraft 51 is designed as a truss structure whichprovides the load path continuity. The same structural elements alsoform the primary structure of the spacecraft 51 and support all majorcomponents such as the perigee motor, the propellant tanks and theremaining payload. Loads induced during launch or landing of the shuttle52 are transmitted through this structure to the cradle 52 and then tothe shuttle 50. Consequently, the structure of the spacecraft 51 servesa dual purpose in that it both stiffens the cradle 52 and provides themajor load paths for the spacecraft 51 itself. The cradle 52 providesmaximum capability for attaching a payload anywhere in the bay of theshuttle 50, and it is well-suited to all types of satellites such asspinners, gyrostats, and three axis spacecraft. Because the cradle 52utilizes the structure of the spacecraft 51 to provide the requiredstiffness, the cradle 52 has a low weight and occupies a low volume,thus increasing available space for the spacecraft 51.

As may be seen in FIGS. 7 and 8, the cradle 52 is made of many separatepieces of relatively thin-walled metal. FIG. 7 is a perspective view oftwo segments of one side of the cradle 52, showing bulkhead, skin andsides. FIG. 8 is an enlarged perspective view of one of the bulkheads 73of the cradle 52. In effect, the cradle 52 comprises a series ofindividual hollow cells or compartments all joined together and coveredwith a skin. The side walls 78 are provided with hollowed portions 79 toreduce the weight. This structure is light and strong and occupies a lowvolume. At the attachment points for the trunnions 65, 66, 67, a heaviermetal block is fastened to the compartmented structure.

Referring now to FIG. 9, the spacecraft 51 is provided with threetrunnion balls 80 (see also FIG. 11) at the three attachment points. Twoof the balls 80 located at the pivot point 53 and at the keel areprovided with swivel blocks 81, 82, (swivel block 81 is also shown inFIGS. 10 and 12) which are fastened around the balls 80 but are free toturn thereon. This arrangement forms a ball-and-socket joint. The swivelblocks 81, 82 slide into slots 77 and 76, respectively, in the cradle.As will be more fully described hereinafter, the ball 80 at the pivotpoint 53 and the ball 80 at the keel are provided with holes throughwhich a locking mechanism secures them to the cradle 52. The ball 80located at the ejection spring 54 (FIGS. 13 and 14) does not have aswivel block and mounts into a locking mechanism associated with theejection spring 54. That ball 80 is provided with a concave depression(FIG. 16) to receive the end of a push rod 122 from the ejection spring54. The ejection spring 54 employs helical coil springs to drive thepush rod. This structure will be described in greater detailhereinafter.

FIG. 10 shows a spacecraft trunnion ball 80 and its swivel block 81,illustrating how the ball 80 is locked into the cradle 52. The member 96(which is part of the spacecraft structure) extending from the ball 80to the spacecraft 51 is shown broken away. The ball 80 is provided witha hole through which a sliding locking bolt 85 passes. The locking bolt85 is provided with a bullet-shaped nose to allow for misalignmentbetween the cradle 52 and the spacecraft 51. Thus, even if thermaldistortions due to differential thermal expansion or basic mechanicaldefects are present, the locking bolt 85 can lock the spacecraft 51 tothe cradle 52.

A pair of bushings 81a are provided in the cradle 52 to receive thelocking bolt 85. The bushings are provided with tapered or wedged-shapedheads which protrude into the slot 77 (FIG. 12) in the cradle 52. Theswivel block 81 is provided with a corresponding taper or wedge shape.The openings in the bushings are sized for a close fit to the lockingbolt 85, whereas the opening in the ball 80 is oversize to allow formovement and swiveling of the swivel block 81 around the ball 80. Thislocking arrangement is provided at both the pivot point 53 in slot 77,and at the keel in slot 76.

At ejection, the spacecraft 51 applies a large downward force at thepivot point 53. This is illustrated in FIG. 11. The ball 80 at thebottom or keel attachment point leaves the cradle 52 at an angle ofapproximately 45 degrees.

FIG. 12 is a perspective view of the locking mechanism which drives thelocking bolt 85. A pair of gearhead DC motors 100, 101 drive through adifferential gear box 102 to a pinion gear 103 which engages rack teeth99 provided along the bottom of the locking bolt 85. The locking bolt 85is provided with a protrusion 104 which engages limit switches 105, 106,107, 108 that control the extent of travel of the locking bolt 85. Thelimit switches 105, 106, 107, 108 are provided in pairs for reliabilityby way of redundancy. The locking bolt 85 is provided with a slot 97along the top thereof which is engaged by a member rigidly fastened tothe locking assembly. The member and slot 97 prevent the locking bolt 85from rotating, and thus disengaging the pinion gear 103 from the rackteeth 99. This arrangement also limits the travel of the locking bolt85. The differential gear box 102 allows either or both of the motors100, 101 to insert or retract the locking bolt 85. Should one of themotors 100, 101 fail, the high drivetrain gearing allows the other oneof the motors 100, 101 to drive the locking bolt 85 with the same axialforce but at one-half the speed of both motors 100, 101.

The locking mechanism has the capability to relock the spacecraftattachments following an unsuccessful deployment attempt if one of therelease mechanisms fails to operate. Because the two permanent magnet DCmotors 100, 101 are reversible, the locking bolt 85 may be inserted orwithdrawn. The purpose of having the capability to relatch thespacecraft 51 to the shuttle 50 is to allow the shuttle 50 to landsafely while bringing back the spacecraft 51 if the deployment isunsuccessful.

One advantage of this arrangement is that remote relatch capabilityprecludes the requirement for extra vehicular activity, where anastronaut would have to enter the shuttle bay following an unsuccessfulejection attempt to perform one of the following operations: (1)manually relatch the spacecraft 51 to the shuttle 50 for safe return ofthe spacecraft 51 intact; (2) manually unlock all latches to allow thespacecraft 51 to be released for normal deployment or to be jettisonedto allow safe return of the shuttle 50. The deployment sequence allowsfor the locking bolt 85 to be retracted at all but the last spacecraftattachment point. Subsequently, a pyrotechnic device is fired, releasingthe spacecraft 51. The sequence is such that if any failure occurs atany release point (including the pyrotechnic release point) allpreceding latches can be relatched, thus securing the spacecraft 51 tothe shuttle 50.

FIGS. 13, 14, 15, 16 show the latching mechanism used at the point wherethe ejection spring 54 applies its force. FIGS. 13 and 14 areperspective views and FIGS. 15 and 16 are side elevations taken at 90degrees with respect to each other. FIG. 13 shows the mechanism in thelocked position, while FIG. 14 shows the mechanism in the released orunlocked position. FIGS. 15 and 16 show the mechanism in the lockedposition. The trunnion ball 80 is rigidly fixed to the spacecraft 51 bythe connecting member 96. The ball 80 is seated in a hemispherical seat112 provided in the cradle 52, best seen in FIG. 16. A locking lever 113is provided with a hemispherical hollow 114 which is clamped over theball 80 locking it into the hemispherical seat 112. The combination ofthe hemispherical hollow 114, the hemispherical seat 112 and the ball 80forms a ball-and-socket joint. The locking lever 113 has a latchingmember 115 which extends or protrudes therefrom and is engaged by abellcrank 116 when the mechanism is in the locked or latched position.The bellcrank 116 is held in the locked position by a bolt 117 which isheld by a separation nut 118. A pushrod 122 from the ejection spring 54engages the bottom of the ball 80, a cavity or depression being providedtherein for engagement with the end of the pushrod 122. The ejectionspring 54 is cocked and exerts a force against the bottom of the ball80.

The explosive separation nut 118 is of a well known or conventional typein which a nut in three sections is held together by a mechanism whichis disengaged by the electrical firing of explosive squibs 123, 124. Apair of squibs 123, 124 is provided for reliability by way ofredundancy. The firing of the explosive squibs 123, 124 causes the nutto separate into its three sections permitting the bolt 117 to fly outof the separation nut 118. The bolt 117 is caught in a hollow boltcatcher 125 which contains spring fingers to engage the head of the bolt117, holding it captive. Upon release of the bolt 117, the bellcrank 116is free to turn and no longer locks the latching member 115 of thelocking lever 113. The force of the ejection spring 54 drives thepushrod 122 upward against the ball 80 causing the spacecraft 51 topivot, resulting in ejection therof.

FIG. 17 shows the locking mechanism for the ejection spring 54 mountedin the end of the cradle 52 and protruding into the slot 75 provided forthe spacecraft trunnion ball 80. The pushrod 122 extends upward from theejection spring 54, which is also mounted inside the cradle 52, but isnot seen in this view.

Referring now to FIG. 18, there is shown the basic skeletal structure ofthe spacecraft 51. As was pointed out hereinbefore, the integratedcombination of the spacecraft 51 and the cradle 52 together provide therequired stiffness in order to support a large diameter spacecraft 51using a narrow U-shaped cradle 52. The structure of the spacecraft 51provides a structural tie between the two ends of the U-shaped cradle 52and the keel which greatly increases the overall stiffness of thecombined structure. The structure of the spacecraft 51 is a trussstructure which provides the load path continuity. The same structuralelements also form the primary structure of the spacecraft 51 andsupport all major components. Loads induced during launch or landing ofthe shuttle 50 are transmitted through this structure to the cradle 52and then to the shuttle 50. Thus, the structure of the spacecraft 51serves a dual purpose in that it stiffens the cradle 52 and provides themajor load paths for the spacecraft 51 itself.

As may be seen in FIG. 18, the truss structure is made of a small squareframe 130 fastened within a larger square frame 131. Other frame members132 tie the corners of the small square frame 130 to the center of thespacecraft 51.

Referring now to FIG. 19, it may be seen that in the other dimension,the truss structure of the spacecraft 51 is in the form of a rectangularframe 133 connected by frame members 134 which form a triangular orA-frame structure. It will be recognized that this overall structure isextremely strong and rigid, and that the three balls 80 which form thetrunnions for the spacecraft 51 are rigidly interconnected.

A segment of the cylindrical solar panel 137 which forms the outercircumference of the spacecraft 51 is shown in FIG. 20. The framestructure supports four liquid apogee motor tanks 138. The frame alsosupports four hydrazine tanks 140 for the reaction control system of thesatellite 51. In addition, there are two helium tanks 141. The antennareflector 142 is shown in its folded down or stowed position.

As may be seen in FIGS. 19 and 20, the spacecraft 51 has imbedded withinits envelope a solid-propellant perigee boost motor 143 surrounded byliquid-propellant apogee motor engines and nozzles 144. After theperigee boost motor 143 has been fired, it is ejected from thespacecraft 51. The arrangement of the perigee boost motor 143 imbeddedin the center of the satellite 51 and surrounded by theliquid-propellant apogee motor engines and nozzles 144 is feasible dueto the large diameter payload bay of the shuttle 50. By use of internalapogee and perigee propulsion stages, the required stowage length forthe satellite 51 is minimized. This provides a compact spacecraft 51which incorporates both the perigee and apogee propulsion stages withinthe normal envelope of the spacecraft 51 and in this way, minimizes thestowage length and the launch cost. Accordingly, multiple payloads cannow be more easily stowed in the payload bay of the shuttle 50. Thegeometry and mass characteristics of this satellite 51 make it a stablespinning vehicle during both the perigee and apogee boost phases, thusno active nutation control is required. In conventional spacecraft, theperigee boost stages are attached in tandem to the exterior of thespacecraft and the result is a much longer configuration, and one thatis inherently unstable because of the ratio of roll-to-pitch inertiabeing less than one, until after the perigee stage is separated from thespacecraft.

The basic spacecraft 51 is 14 feet in diameter with a 10-foot longcylindrical solar panel 137 which represents the controlling outerdimensions of the spacecraft 51. The perigee boost motor 143 is imbeddedin the spacecraft with its thrust axis along the spin axis of thespacecraft 51. The aft end of the nozzle of the perigee kick motor 143extends to the same location as the aft end of the solar panel 137, thusthe perigee kick motor 143 adds no additional length to the spacecraft51. The perigee kick motor 143 is attached to the main spacecraft frameby means of a conical structure.

Surrounding the perigee kick motor 143 are eight propellant tanks 138,140, 141, four for the liquid apogee motor system, and four for thereaction control system which provides spin-up, reorientation andattitude control. The liquid apogee motor tanks 138 are larger in sizeand alternate in location with the hydrazine tanks 140 for the reactioncontrol system. The liquid apogee motor system includes two engineswhich are mounted aft and on each side of the perigee kick motor 143.Two engines are employed to minimize spacecraft coning during thrusting.

The communications payload of the spacecraft 51 is mounted on a despunplatform which is mechanically locked until synchronous orbit isachieved. The despin bearing assembly is located forward of the top endof the perigee kick motor 143.

The basic layout of equipment is designed to produce a ratio of roll topitch inertia greater than one to yield a stable, spinning spacecraft51. The imbedding of the perigee boost motor 143 is a major factor inachieving this goal since it represents almost two-thirds of the totalweight. In the arrangement shown, the center of gravity of the perigeeboost motor 143 is located nearly coincident with the overall center ofgravity, thus greatly minimizing its contribution to the pitch inertia.

The design of the spacecraft 51 allows for separation of the perigeeboost motor 143 and its support structure at the completion of the burnof the perigee boost motor 143. This separation is accomplished by meansof pyrotechnic devices at the interfaces to the main structure.Separation of the burned-out perigee boost motor translates intoadditional payload weight which can be boosted to synchronous orbit bythe liquid-propellant motor propulsion system.

FIGS. 21 and 22 show a spin activated safety circuit for the spacecraft51. If explosive or hazardous pyrotechnic devices were accidentallyarmed and fired while the spacecraft 51 is in the bay of the shuttle 50,the results could be catastrophic. A safety circuit has therefore beenprovided which prevents the arming and firing of these devices until thespacecraft 51 is ejected from the shuttle 50 and spins up to its normalperigee motor firing speed. Conventionally, a spacecraft employsseparation switches to activate the enabling circuitry which arms andfires the motor. If a spacecraft should happen to be ejected onlypartially from the shuttle 50 and then become jammed, those separationswitches could possibly close and activate the enabling circuitrycausing it to arm and fire the motor.

Referring to FIG. 22, the spacecraft main power bus 150 is connectedthrough three separation switches, 151, 152 and 153 which detectclearance between the spacecraft 51 and the cradle 52 which holds it inthe shuttle 50. When an electrical path is completed through theseswitches 151, 152 and 153, power is applied to the bus 154 which isconnected to timers that activate various spacecraft functions,including the firing of spin-up jets to impart rotational velocity tothe spacecraft 51.

Two rotational acceleration sensors known as g-switches 155, 156 aremounted on opposite sides of the spacecraft 51 to sense the fact thatthe spacecraft 51 has indeed spun-up, and therefore must have separatedfrom the shuttle 50. Only when a predetermined rotational speed isreached will the two g-switches 155, 156 close, allowing power to passto the bus 157 which leads to the critical motor arm and fire circuits.Should any malfunction prior to this time cause arm or fire commands tobe issued, they will be ignored. In this way, accidental prematureignition of the perigee kick motor 143 by electronic failures andignition by the sequencer in the shuttle bay if the spacecraft shouldbecome jammed during ejection are both positively precluded. The sensingaxes of the two g-switches 155 and 156 are oriented in opposingdirections, thus the possibility of both switches 155, 156 closingsimultaneously due to linear acceleration of the shuttle 50 isprecluded.

In preparation for launch, the spacecraft 51 is placed in the cradle 52by engaging the three satellite trunnion balls 80 in the slots 75, 76,77 in the cradle 52. Two of the balls 80, the one at the pivot point 53and the other at the keel point, are then locked in place by means ofthe locking bolts 85 driven by the DC motors 100,101. The third ball 80at the ejection spring 54 is locked in place by engaging the lockinglever 113 and the latching member 115 with the bell crank 116, and thenfastening the bolt 117 into the separation nut 118. After the ball 80 isfirmly locked in place, then the ejection spring 54 is cocked. With thesatellite 51 firmly locked into the cradle 52, the assembly of thecradle 52 and satellite 51 is placed into the space shuttle 50 and thecradle trunnions 65, 66, 67 are locked by means of the fittings 70, 71,72 so that the cradle 52 and satellite 51 are firmly mounted to thespace shuttle 50.

After the shuttle 50 has reached the parking orbit 57 (FIG. 4), when itis desired to eject the spacecraft 51, first the locking bolt 85 at thepivot point 53 is retracted by operating the DC motors 100, 101, thenthe locking bolt 85 at the keel point is retracted and finally theseparation nut 118 is fired which allows the ejection spring 54 tosimultaneously impart translation and rotation to the spacecraft 51 forlaunch. The sequence is such that if any failure occurs at any releasepoint, including the pyrotechnic release, all preceding locking bolts 85can be relatched, thus securing the spacecraft 51 to the shuttle 50.

Thus, there has been described a spacecraft and mounting structurespecifically designed to take advantage of launch from the payload bayof the NASA Space Shuttle. It is arranged for gyroscopic deployment withboth spin and translational movement, has a three-point attachment tothe shuttle and a three-point attachment between the spacecraft and thecradle, employs an integrated satellite and cradle structure in whichthe satellite provides strength to the cradle so that a smaller, lightercradle may be used, and it has a locking mechanism for locking thespacecraft to the cradle which may be relocked if circumstances sodictate. The spacecraft employs a solid perigee kick motor internal tothe spacecraft surrounded by the liquid apogee motor so as to provide aminimized stowage length and greater stability because of the ratio ofroll to pitch inertia being less than one, and the satellite is providedwith a spin-activated safety circuit which senses spacecraft spin-up,and therefore separation from the shuttle, before the arming ofhazardous circuits.

It is to be understood that the above-described embodiment of theinvention is merely illustrative of the many possible specificembodiments which represent applications of the principles of thepresent invention. Numerous and varied other arrangements can readily bedevised in accordance with these principles by those skilled in the artwithout departing from the spirit and scope of the invention. Forexample, spring 54 could be replaced by a piston or tangential cablearrangement.

What is claimed is:
 1. An apparatus for gyroscopic ejection of a spacecraft having first and second trunnions on opposite sides thereof, said apparatus comprising:a U-shaped cradle having first and second mating surfaces on opposite sides thereof for mating with said first and second spacecraft trunnions respectively; locking means for normally restraining said first trunnion against movement away from said first mating surface but for selectively releasing said first trunnion so as to allow movement thereof towards the open end of said U-shaped cradle; means for applying a force to said first trunnion, which force is directed towards the open end of said U-shaped cradle; and said second mating surface comprises a slot formed in the side of said cradle, said second trunnion has a ball shaped outer end and a swivel block which is adapted for insertion into said slot and which is fastened to said ball shaped portion of said second trunnion so as to be free to move thereon; whereby when said first trunnion is released by said locking means, a spacecraft which is mounted by said trunnions to said cradle rotates about said ball shaped portion of said second trunnion so as to produce translatory and rotary movement of said spacecraft as it is ejected from said cradle.
 2. An apparatus for gyroscopic ejection of a spacecraft having first and second trunnions on opposite sides thereof, said apparatus comprising:a U-shaped cradle having first and second mating surfaces on opposite sides thereof for mating with said first and second spacecraft trunnions respectively; locking means for normally restraining said first trunnion against movement away from said first mating surface but for selectively releasing said first trunnion so as to allow movement thereof towards the open end of said U-shaped cradle; means for applying a force to said first trunnion, which force is directed towards the open end of said U-shaped cradle; a third trunnion on said spacecraft between said first and second trunnions; and means forming a third mating surface on said cradle for slidably receiving said third trunnion; whereby when said first trunnion is released by said locking means, a spacecraft which is mounted by means of its trunnions to said cradle is rotated at said second trunnion as said third trunnion slides at an angle from said third mating surface, to produce rotary and translatory movement of said spacecraft as it is ejected from said cradle.
 3. An apparatus for gyroscopic ejection of a spacecraft having first and second trunnions on opposite sides thereof, said apparatus comprising:a U-shaped cradle having first and second mating surfaces on opposite sides thereof for mating with said first and second spacecraft trunnings respectively; locking means for normally restraining said first trunnion against movement away from said first mating surface but for selectively releasing said first trunnion so as to allow movement thereof towards the open end of said U-shaped cradle; means for applying a force to said first trunnion, which force is directed towards the open end of said U-shaped cradle; said second mating surface comprising a slot formed in the side of said cradle, said second trunnion having a ball shaped outer end and a swivel block fitting into said slot and which is fastened to said ball shaped outer end of said second trunnion so as to be free to move thereon; a third trunnion on said spacecraft between said first and second trunnions; and means forming a guide slot in said cradle for receiving said third trunnion when said first and second trunnions are engaged with said first and second mating surfaces, respectively, said third trunnion having a ball shaped outer end and a swivel thereon fitting into said guide slot; whereby when said first trunnion is released by said locking means, a spacecraft which is mounted by means of its trunnions to said cradle is rotated about said ball shaped outer end at said second trunnion, as said third trunnion slides from said guide slots, to produce rotary and translatory movement of said spacecraft as it is ejected from said cradle. 